US20210148241A1 - Vane retention feature - Google Patents
Vane retention feature Download PDFInfo
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- US20210148241A1 US20210148241A1 US16/689,529 US201916689529A US2021148241A1 US 20210148241 A1 US20210148241 A1 US 20210148241A1 US 201916689529 A US201916689529 A US 201916689529A US 2021148241 A1 US2021148241 A1 US 2021148241A1
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- Prior art keywords
- lip
- platform
- airfoil
- support ring
- sloped surface
- Prior art date
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/042—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector fixing blades to stators
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
- F01D5/284—Selection of ceramic materials
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
- F01D25/246—Fastening of diaphragms or stator-rings
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
- F01D5/282—Selecting composite materials, e.g. blades with reinforcing filaments
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2230/00—Manufacture
- F05D2230/60—Assembly methods
- F05D2230/64—Assembly methods using positioning or alignment devices for aligning or centring, e.g. pins
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/10—Stators
- F05D2240/12—Fluid guiding means, e.g. vanes
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/80—Platforms for stationary or moving blades
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2250/00—Geometry
- F05D2250/30—Arrangement of components
- F05D2250/38—Arrangement of components angled, e.g. sweep angle
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/30—Retaining components in desired mutual position
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/30—Retaining components in desired mutual position
- F05D2260/36—Retaining components in desired mutual position by a form fit connection, e.g. by interlocking
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2300/00—Materials; Properties thereof
- F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
- F05D2300/603—Composites; e.g. fibre-reinforced
- F05D2300/6033—Ceramic matrix composites [CMC]
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- a gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-energy exhaust gas flow. The high-energy exhaust gas flow expands through the turbine section to drive the compressor and the fan section.
- the compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.
- any of the fan section, the turbine section, and the compressor section include airfoils, such as for fan, compressor, or turbine blades.
- Baffles are known and used in cooled gas turbine engine airfoils, such as turbine vanes.
- a baffle is situated in a cavity in the airfoil and serves to distribute cooling air to precise locations in the airfoil.
- An airfoil assembly includes an airfoil.
- the airfoil has first and second platforms and an airfoil between the first and second platforms.
- a support ring is configured to retain the first platform.
- the support ring has a lip which extends radially inward from the support ring.
- the lip is configured to engage an axial face of the first platform.
- the lip has a primary retention feature and a secondary retention feature. The primary and secondary retention features are configured to retain the first face of the first platform with respect to the annular ring.
- the primary retention feature includes a sloped surface on the lip.
- the sloped surface is configured to mate with a sloped surface on a face of the first platform.
- the sloped surface on the lip is sloped with respect to a radial direction of the support ring.
- the secondary retention feature includes a secondary lip which extends from a radially inward end of the sloped surface of the lip.
- the first platform is a radially outer platform.
- the first platform includes a plurality of ceramic matrix composite plies is arranged perpendicular to a radial direction of the airfoil.
- a second support ring is configured to retain the second platform.
- the lip is a first lip.
- the first lip engages a first axial face of the first platform.
- the support ring further includes a second lip which is configured to engage a second axial face of the first platform.
- the first axial face is an aft face and the second axial face is a forward face.
- a location feature is configured to radially locate the airfoil with respect to the support ring.
- a support structure for an airfoil includes an annular ring.
- the annular ring has a first axial side and a second axial side opposite the first axial side.
- the first and second lips extend radially towards the center of the annular ring from the first and second axial sides respectively.
- the first lip has a primary retention feature and a secondary retention feature.
- the primary and secondary retention features are configured to retain a platform of an airfoil with respect to the annular ring.
- the primary retention feature includes a sloped surface on the first lip.
- the sloped surface is configured to mate with a sloped surface on a face of the airfoil platform.
- the sloped surface is sloped with respect to a radial direction of the annular ring.
- the secondary retention feature includes a secondary lip which extends from a radially inward end of the sloped surface of the first lip.
- the support ring includes a groove which is configured to receive a tongue on the platform to locate the airfoil with respect to the support ring.
- a method of retaining an airfoil assembly includes retaining an airfoil platform with respect to a support ring via a lip, which extends radially inward from the support ring.
- the lip is configured to engage an axial face of the platform.
- the lip has a primary retention feature and a secondary retention feature. The primary and secondary retention features are configured to retain the axial face of the platform with respect to the support ring.
- the primary retention feature includes a sloped surface on the first lip.
- the sloped surface is configured to mate with a sloped surface on a face of the airfoil platform.
- the sloped surface is sloped with respect to a radial direction of the support ring.
- the secondary retention feature includes a secondary lip which extends from a radially inward end of the sloped surface of the lip.
- the airfoil platform is located with respect to the support thing by locating a tongue on the airfoil platform in a groove of the support ring.
- FIG. 1 is a schematic view of an example gas turbine engine.
- FIG. 2 is a schematic view of an airfoil for the example gas turbine engine of FIG. 1 .
- FIG. 3A is a schematic view of an airfoil for the example gas turbine engine of FIG. 1 in a support structure.
- FIG. 3B is a detail view of the airfoil and support structure of FIG. 3A .
- FIG. 1 schematically illustrates a gas turbine engine 20 .
- the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22 , a compressor section 24 , a combustor section 26 and a turbine section 28 .
- the fan section 22 drives air along a bypass flow path B in a bypass duct defined within a nacelle 15 , and also drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28 .
- the exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38 .
- Terms such as “axial,” “radial,” “circumferential,” and variations of these terms are made with reference to the engine central axis A. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.
- the low speed spool 30 generally includes an inner shaft 40 that interconnects, a first (or low) pressure compressor 44 and a first (or low) pressure turbine 46 .
- the inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive a fan 42 at a lower speed than the low speed spool 30 .
- the high speed spool 32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor 52 and a second (or high) pressure turbine 54 .
- a combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54 .
- a mid-turbine frame 57 of the engine static structure 36 may be arranged generally between the high pressure turbine 54 and the low pressure turbine 46 .
- the mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28 .
- the inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
- the core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52 , mixed and burned with fuel in the combustor 56 , then expanded over the high pressure turbine 54 and low pressure turbine 46 .
- the mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C.
- the turbines 46 , 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22 , compressor section 24 , combustor section 26 , turbine section 28 , and fan drive gear system 48 may be varied.
- gear system 48 may be located aft of the low pressure compressor, or aft of the combustor section 26 or even aft of turbine section 28 , and fan 42 may be positioned forward or aft of the location of gear system 48 .
- the engine 20 in one example is a high-bypass geared aircraft engine.
- the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10)
- the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3
- the low pressure turbine 46 has a pressure ratio that is greater than about five.
- the engine 20 bypass ratio is greater than about ten (10:1)
- the fan diameter is significantly larger than that of the low pressure compressor 44
- the low pressure turbine 46 has a pressure ratio that is greater than about five 5:1.
- Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
- the geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1 and less than about 5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
- the fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters).
- TSFC Thrust Specific Fuel Consumption
- Low fan pressure ratio is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system.
- the low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45.
- Low corrected fan tip speed is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7°R)] ⁇ circumflex over ( ) ⁇ 0.5.
- the “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 meters/second).
- FIG. 2 schematically shows an airfoil 100 from the turbine section 28 of the engine 20 .
- a plurality of airfoils 100 are situated in a circumferential row about the engine central axis A.
- the airfoil 100 includes a first or inner platform 102 , a second or outer platform 104 , and an airfoil section 106 that spans between the inner and outer platforms 102 / 104 .
- FIG. 3 shows a cross-sectional view of the airfoil section 106 along the section line A-A in FIG. 2 .
- the airfoil section 106 includes an airfoil outer wall 108 that delimits the profile of the airfoil section 106 .
- the outer wall 108 defines a leading end 108 a , a trailing end 108 b , and first and second sides 108 c / 108 d ( FIG. 3 ) that join the leading and trailing ends 108 a / 108 b .
- the first side 108 c is a pressure side
- the second side 108 d is a suction side.
- the outer wall 108 circumscribes an internal cavity 110 .
- the cavity 110 may be a single cavity or a sub-cavity, for example.
- the cavity 110 can be configured to receive a baffle or spar (not shown).
- the airfoil 100 is formed of a ceramic material, such as a ceramic matrix composite (CMC) material.
- the CMC includes a ceramic matrix and ceramic fibers disposed in the ceramic matrix.
- the ceramic matrix may be, but is not limited to, silicon carbide (SiC) and the ceramic fibers may be, but are not limited to, silicon carbide (SiC) fibers.
- the CMC is comprised of fiber plies that are arranged in a stacked configuration and formed to the desired geometry of the airfoil 100 .
- the fiber plies may be layers or tapes that are laid-up one on top of the other to form the stacked configuration.
- the fiber plies may be woven or unidirectional, for example.
- the airfoil 100 is a continuous body in that the fiber plies are uninterrupted through the platforms 102 / 104 and the airfoil section 106 .
- the airfoil 100 may also be a monolithic ceramic material, such as a silicon-containing ceramic. Examples of such ceramics include silicon nitride and silicon carbide.
- the airfoil 100 is shown in the turbine section 28 (though the airfoil 100 could be in other section in other examples, as discussed above).
- the platforms 102 / 104 are supported in inner and outer annular support rings 112 / 114 , respectively.
- the inner and outer support rings 112 / 114 are connected to static structures in the engine 20 .
- the platforms 102 / 104 each have a forward axial face 102 a / 104 a and an aft axial face 104 a / 104 b . As best shown in FIG.
- the outer support ring 114 includes a retention feature in the form of lips 116 which engages the forward face 104 a and aft face 104 b of the outer platform 104 .
- the lips 116 extend from the outer support ring 114 in a radially inward direction to capture the outer platform 104 .
- the outer platform 104 is comprised of plies 200 of CMC material, as discussed above. Interlaminar regions 202 are defined between adjacent plies.
- the outer platform 104 /outer support ring 114 include a location feature 118 which locates the outer platform 104 with respect to the outer support ring 114 .
- the location feature 118 can include a shim which is configured to radially locate the vane, in some examples.
- the outer platform 104 includes a primary and secondary retention feature.
- the primary retention features includes a pair of mating sloped surfaces 122 a / 122 b .
- the mating sloped surface 122 a / 122 b are angled with respect to the radial direction R by an angle ⁇ .
- the first sloped surface 122 a is on the aft end 104 b of the outer platform 104 .
- the second sloped surface 122 b is on the lip 116 .
- the outer platform 104 experiences loads L perpendicular to the radial direction R as shown in FIG. 3B .
- the sloped surfaces 122 a / 122 b counteract the loads L to retain the outer platform 104 with respect to the outer support ring 114 .
- the secondary retention feature is a secondary lip 124 extending radially inward from the lip 116 .
- the secondary lip 124 extends from a radially inward end of the mating surface 122 b .
- the secondary lip 124 is situated to engage a radial surface 126 of the forward edge 104 b of the platform 104 which is radially inward from the sloped surface 122 a in the event that the primary retention feature 122 a / 122 b is lost or otherwise becomes inoperable.
- the secondary lip 124 extends substantially to a radially innermost extend of the outer platform 104 .
- the outer platform 104 may experience a shear force S opposite the load L shown in FIG. 3B .
- the plies 200 are generally arranged parallel to the shear force S and load L.
- the shear force S could contribute to interlaminar stresses at interlaminar regions 202 between individual plies 200 of the CMC material of the platform 104 .
- the shear force S is higher in some radial areas of the outer platform 104 as compared to other radial areas, e.g., in the event of a partial loss of the primary retention feature 122 a / 122 b , were the load L and thus shear force S are not evenly distributed to adjacent plies 200 .
- These interlaminar stresses if large enough, could have the potential to delaminate the plies 200 from one another, which in turn could lead to damage in the outer platform 104 .
- the secondary lip 124 prevents the airfoil 100 from liberating from the outer support ring 114 due to load L in the event of full failure of the primary retention feature 122 a / 122 b.
Abstract
Description
- A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-energy exhaust gas flow. The high-energy exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.
- Any of the fan section, the turbine section, and the compressor section include airfoils, such as for fan, compressor, or turbine blades. Baffles are known and used in cooled gas turbine engine airfoils, such as turbine vanes. Typically, a baffle is situated in a cavity in the airfoil and serves to distribute cooling air to precise locations in the airfoil.
- An airfoil assembly according to an exemplary embodiment of this disclosure, among other possible things includes an airfoil. The airfoil has first and second platforms and an airfoil between the first and second platforms. A support ring is configured to retain the first platform. The support ring has a lip which extends radially inward from the support ring. The lip is configured to engage an axial face of the first platform. The lip has a primary retention feature and a secondary retention feature. The primary and secondary retention features are configured to retain the first face of the first platform with respect to the annular ring.
- In a further example of the foregoing, the primary retention feature includes a sloped surface on the lip. The sloped surface is configured to mate with a sloped surface on a face of the first platform.
- In a further example of any of the foregoing, the sloped surface on the lip is sloped with respect to a radial direction of the support ring.
- In a further example of any of the foregoing, the secondary retention feature includes a secondary lip which extends from a radially inward end of the sloped surface of the lip.
- In a further example of any of the foregoing, the first platform is a radially outer platform.
- In a further example of any of the foregoing, the first platform includes a plurality of ceramic matrix composite plies is arranged perpendicular to a radial direction of the airfoil.
- In a further example of any of the foregoing, a second support ring is configured to retain the second platform.
- In a further example of any of the foregoing, the lip is a first lip. The first lip engages a first axial face of the first platform. The support ring further includes a second lip which is configured to engage a second axial face of the first platform.
- In a further example of any of the foregoing, the first axial face is an aft face and the second axial face is a forward face.
- In a further example of any of the foregoing, a location feature is configured to radially locate the airfoil with respect to the support ring.
- A support structure for an airfoil according to an exemplary embodiment of this disclosure, among other possible things includes an annular ring. The annular ring has a first axial side and a second axial side opposite the first axial side. The first and second lips extend radially towards the center of the annular ring from the first and second axial sides respectively. The first lip has a primary retention feature and a secondary retention feature. The primary and secondary retention features are configured to retain a platform of an airfoil with respect to the annular ring.
- In a further example of the foregoing, the primary retention feature includes a sloped surface on the first lip. The sloped surface is configured to mate with a sloped surface on a face of the airfoil platform.
- In a further example of any of the foregoing, the sloped surface is sloped with respect to a radial direction of the annular ring.
- In a further example of any of the foregoing, the secondary retention feature includes a secondary lip which extends from a radially inward end of the sloped surface of the first lip.
- In a further example of any of the foregoing, the support ring includes a groove which is configured to receive a tongue on the platform to locate the airfoil with respect to the support ring.
- A method of retaining an airfoil assembly according to an exemplary embodiment of this disclosure, among other possible things includes retaining an airfoil platform with respect to a support ring via a lip, which extends radially inward from the support ring. The lip is configured to engage an axial face of the platform. The lip has a primary retention feature and a secondary retention feature. The primary and secondary retention features are configured to retain the axial face of the platform with respect to the support ring.
- In a further example of the foregoing, the primary retention feature includes a sloped surface on the first lip. The sloped surface is configured to mate with a sloped surface on a face of the airfoil platform.
- In a further example of any of the foregoing, the sloped surface is sloped with respect to a radial direction of the support ring.
- In a further example of any of the foregoing, the secondary retention feature includes a secondary lip which extends from a radially inward end of the sloped surface of the lip.
- In a further example of any of the foregoing, the airfoil platform is located with respect to the support thing by locating a tongue on the airfoil platform in a groove of the support ring.
- Although the different examples have the specific components shown in the illustrations, embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples.
- These and other features disclosed herein can be best understood from the following specification and drawings, the following of which is a brief description.
-
FIG. 1 is a schematic view of an example gas turbine engine. -
FIG. 2 is a schematic view of an airfoil for the example gas turbine engine ofFIG. 1 . -
FIG. 3A is a schematic view of an airfoil for the example gas turbine engine ofFIG. 1 in a support structure. -
FIG. 3B is a detail view of the airfoil and support structure ofFIG. 3A . -
FIG. 1 schematically illustrates agas turbine engine 20. Thegas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates afan section 22, acompressor section 24, acombustor section 26 and aturbine section 28. Thefan section 22 drives air along a bypass flow path B in a bypass duct defined within anacelle 15, and also drives air along a core flow path C for compression and communication into thecombustor section 26 then expansion through theturbine section 28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. - The
exemplary engine 20 generally includes alow speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an enginestatic structure 36 viaseveral bearing systems 38. Terms such as “axial,” “radial,” “circumferential,” and variations of these terms are made with reference to the engine central axis A. It should be understood thatvarious bearing systems 38 at various locations may alternatively or additionally be provided, and the location ofbearing systems 38 may be varied as appropriate to the application. - The
low speed spool 30 generally includes aninner shaft 40 that interconnects, a first (or low)pressure compressor 44 and a first (or low)pressure turbine 46. Theinner shaft 40 is connected to thefan 42 through a speed change mechanism, which in exemplarygas turbine engine 20 is illustrated as a gearedarchitecture 48 to drive afan 42 at a lower speed than thelow speed spool 30. Thehigh speed spool 32 includes anouter shaft 50 that interconnects a second (or high)pressure compressor 52 and a second (or high)pressure turbine 54. Acombustor 56 is arranged inexemplary gas turbine 20 between thehigh pressure compressor 52 and thehigh pressure turbine 54. A mid-turbine frame 57 of the enginestatic structure 36 may be arranged generally between thehigh pressure turbine 54 and thelow pressure turbine 46. The mid-turbine frame 57 furthersupports bearing systems 38 in theturbine section 28. Theinner shaft 40 and theouter shaft 50 are concentric and rotate via bearingsystems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes. - The core airflow is compressed by the
low pressure compressor 44 then thehigh pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over thehigh pressure turbine 54 andlow pressure turbine 46. The mid-turbine frame 57 includesairfoils 59 which are in the core airflow path C. Theturbines low speed spool 30 andhigh speed spool 32 in response to the expansion. It will be appreciated that each of the positions of thefan section 22,compressor section 24,combustor section 26,turbine section 28, and fandrive gear system 48 may be varied. For example,gear system 48 may be located aft of the low pressure compressor, or aft of thecombustor section 26 or even aft ofturbine section 28, andfan 42 may be positioned forward or aft of the location ofgear system 48. - The
engine 20 in one example is a high-bypass geared aircraft engine. In a further example, theengine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the gearedarchitecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and thelow pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, theengine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of thelow pressure compressor 44, and thelow pressure turbine 46 has a pressure ratio that is greater than about five 5:1.Low pressure turbine 46 pressure ratio is pressure measured prior to inlet oflow pressure turbine 46 as related to the pressure at the outlet of thelow pressure turbine 46 prior to an exhaust nozzle. The gearedarchitecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1 and less than about 5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. - A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The
fan section 22 of theengine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7°R)]{circumflex over ( )}0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 meters/second). -
FIG. 2 schematically shows anairfoil 100 from theturbine section 28 of theengine 20. A plurality ofairfoils 100 are situated in a circumferential row about the engine central axis A. Theairfoil 100 includes a first orinner platform 102, a second orouter platform 104, and anairfoil section 106 that spans between the inner andouter platforms 102/104.FIG. 3 shows a cross-sectional view of theairfoil section 106 along the section line A-A inFIG. 2 . Theairfoil section 106 includes an airfoilouter wall 108 that delimits the profile of theairfoil section 106. Theouter wall 108 defines aleading end 108 a, a trailingend 108 b, and first andsecond sides 108 c/108 d (FIG. 3 ) that join the leading and trailing ends 108 a/108 b. In this example, thefirst side 108 c is a pressure side and thesecond side 108 d is a suction side. Theouter wall 108 circumscribes aninternal cavity 110. Thecavity 110 may be a single cavity or a sub-cavity, for example. Thecavity 110 can be configured to receive a baffle or spar (not shown). - In one example, the
airfoil 100 is formed of a ceramic material, such as a ceramic matrix composite (CMC) material. For example, the CMC includes a ceramic matrix and ceramic fibers disposed in the ceramic matrix. The ceramic matrix may be, but is not limited to, silicon carbide (SiC) and the ceramic fibers may be, but are not limited to, silicon carbide (SiC) fibers. The CMC is comprised of fiber plies that are arranged in a stacked configuration and formed to the desired geometry of theairfoil 100. For instance, the fiber plies may be layers or tapes that are laid-up one on top of the other to form the stacked configuration. The fiber plies may be woven or unidirectional, for example. At least a portion of the fiber plies are continuous through theplatforms 102/104 and theairfoil section 106. In this regard, theairfoil 100 is a continuous body in that the fiber plies are uninterrupted through theplatforms 102/104 and theairfoil section 106. Theairfoil 100 may also be a monolithic ceramic material, such as a silicon-containing ceramic. Examples of such ceramics include silicon nitride and silicon carbide. - Referring to
FIGS. 3A-B , theairfoil 100 is shown in the turbine section 28 (though theairfoil 100 could be in other section in other examples, as discussed above). Theplatforms 102/104 are supported in inner and outer annular support rings 112/114, respectively. The inner and outer support rings 112/114 are connected to static structures in theengine 20. Theplatforms 102/104 each have a forwardaxial face 102 a/104 a and an aftaxial face 104 a/104 b. As best shown inFIG. 3A , theouter support ring 114 includes a retention feature in the form oflips 116 which engages theforward face 104 a andaft face 104 b of theouter platform 104. In general, thelips 116 extend from theouter support ring 114 in a radially inward direction to capture theouter platform 104. - Referring to
FIG. 3B , theouter platform 104 is comprised ofplies 200 of CMC material, as discussed above.Interlaminar regions 202 are defined between adjacent plies. - With continued reference to
FIG. 3B , in one example, theouter platform 104/outer support ring 114 include alocation feature 118 which locates theouter platform 104 with respect to theouter support ring 114. Thelocation feature 118 can include a shim which is configured to radially locate the vane, in some examples. - The
outer platform 104 includes a primary and secondary retention feature. The primary retention features includes a pair of mating slopedsurfaces 122 a/122 b. The mating slopedsurface 122 a/122 b are angled with respect to the radial direction R by an angle α. The firstsloped surface 122 a is on theaft end 104 b of theouter platform 104. The secondsloped surface 122 b is on thelip 116. Theouter platform 104 experiences loads L perpendicular to the radial direction R as shown inFIG. 3B . The sloped surfaces 122 a/122 b counteract the loads L to retain theouter platform 104 with respect to theouter support ring 114. - The secondary retention feature is a
secondary lip 124 extending radially inward from thelip 116. Thesecondary lip 124 extends from a radially inward end of themating surface 122 b. Thesecondary lip 124 is situated to engage aradial surface 126 of theforward edge 104 b of theplatform 104 which is radially inward from the slopedsurface 122 a in the event that the primary retention feature 122 a/122 b is lost or otherwise becomes inoperable. In one example, thesecondary lip 124 extends substantially to a radially innermost extend of theouter platform 104. - In particular, in the event of partial loss of the primary retention feature 122 a/122 b, the
outer platform 104 may experience a shear force S opposite the load L shown inFIG. 3B . Theplies 200 are generally arranged parallel to the shear force S and load L. The shear force S could contribute to interlaminar stresses atinterlaminar regions 202 betweenindividual plies 200 of the CMC material of theplatform 104. This is especially true when the shear force S is higher in some radial areas of theouter platform 104 as compared to other radial areas, e.g., in the event of a partial loss of the primary retention feature 122 a/122 b, were the load L and thus shear force S are not evenly distributed toadjacent plies 200. These interlaminar stresses, if large enough, could have the potential to delaminate theplies 200 from one another, which in turn could lead to damage in theouter platform 104. Also, thesecondary lip 124 prevents theairfoil 100 from liberating from theouter support ring 114 due to load L in the event of full failure of the primary retention feature 122 a/122 b. - Although the different examples are illustrated as having specific components, the examples of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from any of the embodiments in combination with features or components from any of the other embodiments.
- The foregoing description shall be interpreted as illustrative and not in any limiting sense. A worker of ordinary skill in the art would understand that certain modifications could come within the scope of this disclosure. For these reasons, the following claims should be studied to determine the true scope and content of this disclosure.
Claims (25)
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US16/689,529 US11268393B2 (en) | 2019-11-20 | 2019-11-20 | Vane retention feature |
EP20208571.8A EP3825518B1 (en) | 2019-11-20 | 2020-11-19 | Vane retention assembly |
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US16/689,529 US11268393B2 (en) | 2019-11-20 | 2019-11-20 | Vane retention feature |
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US20210148241A1 true US20210148241A1 (en) | 2021-05-20 |
US11268393B2 US11268393B2 (en) | 2022-03-08 |
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EP4290051A1 (en) * | 2022-06-03 | 2023-12-13 | RTX Corporation | Vane arc segment with single-sided platform |
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US4422827A (en) * | 1982-02-18 | 1983-12-27 | United Technologies Corporation | Blade root seal |
US20130052030A1 (en) * | 2011-08-23 | 2013-02-28 | Michael G. McCaffrey | Ceramic matrix composite vane structure with overwrap for a gas turbine engine |
US20160201488A1 (en) * | 2014-10-02 | 2016-07-14 | United Technologies Corporation | Vane assembly with trapped segmented vane structures |
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US3362681A (en) | 1966-08-24 | 1968-01-09 | Gen Electric | Turbine cooling |
US8038389B2 (en) | 2006-01-04 | 2011-10-18 | General Electric Company | Method and apparatus for assembling turbine nozzle assembly |
US7874795B2 (en) | 2006-09-11 | 2011-01-25 | General Electric Company | Turbine nozzle assemblies |
US9343467B2 (en) | 2014-08-28 | 2016-05-17 | Kabushiki Kaisha Toshiba | Semiconductor device |
US10370986B2 (en) | 2015-07-24 | 2019-08-06 | General Electric Company | Nozzle and nozzle assembly for gas turbine engine |
US10370990B2 (en) | 2017-02-23 | 2019-08-06 | General Electric Company | Flow path assembly with pin supported nozzle airfoils |
US10975706B2 (en) | 2019-01-17 | 2021-04-13 | Raytheon Technologies Corporation | Frustic load transmission feature for composite structures |
-
2019
- 2019-11-20 US US16/689,529 patent/US11268393B2/en active Active
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2020
- 2020-11-19 EP EP20208571.8A patent/EP3825518B1/en active Active
Patent Citations (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4422827A (en) * | 1982-02-18 | 1983-12-27 | United Technologies Corporation | Blade root seal |
US20130052030A1 (en) * | 2011-08-23 | 2013-02-28 | Michael G. McCaffrey | Ceramic matrix composite vane structure with overwrap for a gas turbine engine |
US20160201488A1 (en) * | 2014-10-02 | 2016-07-14 | United Technologies Corporation | Vane assembly with trapped segmented vane structures |
Cited By (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP4290051A1 (en) * | 2022-06-03 | 2023-12-13 | RTX Corporation | Vane arc segment with single-sided platform |
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US11268393B2 (en) | 2022-03-08 |
EP3825518B1 (en) | 2022-12-28 |
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